This invention relates generally to gas turbine engines, and more specifically to bleed scoops used in gas turbine engine high-pressure compressors.
At least some known gas turbine engines include a compressor, a combustor, and at least one turbine. The compressor compresses air that is mixed with fuel and channeled to the combustor. The mixture is then ignited for generating hot combustion gases, and the combustion gases are channeled to the turbine which extracts energy from the combustion gases for powering the compressor, as well as producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator.
The compressor includes a rotor assembly and a stator assembly. The rotor assembly includes a plurality of rotor blades extending radially outward from a disk. More specifically, each rotor blade extends radially between a platform adjacent the disk, to a tip. A gas flowpath through the rotor assembly is bound radially inward by the rotor blade platforms, and radially outward by a plurality of shrouds.
The stator assembly includes a plurality of stator vanes that form nozzles that direct the compressed gas entering the compressor to the rotor blades. The stator vanes extend radially between a root platform and an outer band. The outer band retains the stator vanes and mounts the stator assembly within the compressor casing. The stator assembly also forms a bleed plenum between the outer band and the compressor casing.
Within at least some known gas turbine engines, a portion of high-pressure air is extracted or bled from the compressor and directed to a clearance control system. More specifically, a bleed scoop in the stator assembly outer band channels the extracted air to a bleed plenum for the clearance control system. However, the effectiveness of the clearance control system may be reduced because the pressure of the extracted air drops as it passes through the bleed scope. Further, the bleed scoop can provide compressed air with both static and dynamic pressure components.
In an exemplary embodiment, a method for operating a gas turbine engine includes a compressor including stator vanes, a stator assembly, a compressor casing, and a bleed plenum. Adjacent stator vanes define a region between the vanes of high static pressure. The method includes extracting pressurized air from the compressor into at least one bleed scoop formed within the high static pressure regions defined by the adjacent stator vanes and directing the airflow from the bleed scoops into a bleed plenum.
In another aspect, a compressor for a gas turbine engine, the compressor comprises a casing and a stator assembly. The stator assembly comprises a plurality of stator vanes, adjacent stator vanes defining a static high-pressure region between the vanes. The stator assembly also defines at least one bleed scoop extending at least partially through the static high-pressure region adjacent the stator vanes and in flow communication with the gas flowpath and a bleed plenum.
In another aspect, a gas turbine engine includes a compressor including a casing and a stator assembly. The casing extends around the stator assembly, the stator assembly including a plurality of stator vanes, and a plurality of bleed scoops. Adjacent stator vanes define a static high-pressure region, the bleed scoops formed at least partially within the high-pressure region.